Turbine stator-combustor structure

ABSTRACT

A TURBINE STATOR-COMBUSTOR STRUCTURE IN WHICH THE TURBINE STATOR VANES ARE INTEGRATED INTO AN ANNULAR COMBUSTOR RESULTING IN A MUCH SHORTER COMBINATION. THE STATOR VANES ARE CONVECTION AND FILM COOLED WHILE IMPINGEMENT COOLING OF THE LEADING EDGES IS ALSO PROVIDED.

P 23, 1971 J. M. VAUGHT 3,608,310

TURBINE STATORCOMBUSTOR STRUCTURE Filed June 27, 1966 ATTORNEY United States Patent 3,608,310 TURBINE STATOR-COMBUSTOR STRUCTURE John M. Vaught, Indianapolis, Ind., 'assignor to General Motors Corporation, Detroit, Mich. Filed June 27, 1966, Ser. No. 560,658 Int. Cl. F02c 7/12 US. Cl. 60-39.66 3 Claims ABSTRACT OF THE DISCLOSURE A turbine stator-combustor structure in which the turbine stator vanes are integrated into an annular combustor resulting in a much shorter combination. The stator vanes are convection and film cooled while impingement cooling of the leading edges is also provided.

My invention relates generally to combustor equipment for a gas turbine engine or the like and more specifically to a turbine stator-combustor structure which is especially suitable for, but not limited to, use with a lift type gas turbine engine because of its reduced axial length.

For some time now, the feasibility of VTOL aircraft has been under study. One basic concept under study is the inclusion of lift engines vertically mounted in the aircraft which are used during take-off and landing only and are inoperable during normal flight. With the lift engine mounted vertically, it is highly desirable to reduce its axial length as much as practicable. To this end, in my invention, I have integrated a turbine inlet stator into the dilution zone of an annular combustor to decrease the axial length normally required by these components. Ingeneral, the material now available for the turbine stator or nozzle inlet vane row has a temperture limit which is far below the temperatures available from the combustion of liquid hydrocarbon fuels in air at or near the chemically correct fuel/ air ratio (stoichiometric ratio). For this reason, it is necessary to include a section in the combustor in which the combustion gases are diluted with a large excess of secondary air to reduce their temperature before they reach the turbine stator.

The area of the combustor in which dilution or secondary air is admitted is known as the dilution zone which heretofore has added length to the combustor and turbine inlet stator structure. In other words, without a temperature limit problem, no dilution zone would be required and the added length of the dilution zone eliminated. The temperature limit problem of the stator exists, however, and it increases as the distance between the primary combustion zone and the inlet stator decreases. Therefore, any decrease in this distance (and consequently the axial length of the structure) must take into account the temperature limit of the inlet nozzle vane stator. In my unique scheme, I utilize the dilution or secondary air in a unique manner to both decrease the temperature of the combustion gases and to give an impingement cooling effect to the vanes at their leading edges or most critical areas. In addition, air is bled from the secondary air stream to convection cool the vanes themselvesthe cooling air being discharged through the vanes to provide film cooling of their outer surfaces. Because the air bleed for convection cooling flows out of the blade into the main gas stream, another advantage is realized in that the effect to engine performance of bleeding compressed air is minimized. Accordingly, the object of my invention in its broadest aspect is to provide an operable turbine stator-combustor structure for a gas turbine or the like having an axial length shorter than any available comparable structure.

Another object of my invention is to provide a relatively short combustor structure having a turbine stator ice incorporated in its dilution zone, in which the hydrocarbon fuel may be efliciently burned, in which the combustion gases are cooled to a temperature which is capable of being withstood by the turbine stator and in which does not appreciably affect engine performance.

Another object of my invention is to provide turbine stator-combustor structure having a relatively short axial length and including means for convection cooling the turbine stator vanes, film cooling of their outer surfaces, and impingement cooling of their leading edges so that they can withstand the environmental temperature produced by the combustion of a stoichiometric mixture of fuel and air in the combustor.

With these and other objects in view, as will hereinafter more fully appear, and which will be more particularly pointed out in'the appended claims, reference is now made to the following description taken in connection with the accompanying drawings in which:

FIG. 1 is a section taken on a plane containing the longitudinal axis of a portion of a gas turbine engine showing an integrated turbine stator-combustor in accordance with my invention and its relationship to the engine.

FIG. 2 is a section taken along the line 2--2 of FIG. 1 and looking in the direction of the arrows.

Referring now to the drawings and more particularly to FIGS. 1 and 2, the gas turbine engine combustion portion is indicated generally at 12. The outer casing 14 of the engine and a wall 16 mounted within it define an annular air casing which receives compressed air at inlet 17. An annular combustor 18 mounted in the air casing and spaced from it provides outer and inner annular air passages 20 and 22. Mounted on the forward end of the combustor 18 is a deflector 23 which proportions the air flow from inlet 17 between the air passages 19 and the annular air passages 20 and 22. A fuel ring 24 mounted inside the deflector 23 includes a number of circumferentially spaced fuel spray nozzles 26 which protrude into the forward end of the combustor 18.

Mounted radially across the aft end of the combustor 18 are a number of circumferentially spaced nozzle vanes 28. These nozzle vanes are hollow and have open radially outer and inner ends 30 and 32 which are in fluid communication with the annular air passages 20 and 22, respectively. The walls 33 providing the aerodynamic contour of the vanes have a number of apertures 35 of various size, larger ones being provided in the midportions of the vanes. Apertures 37 in the combustor inner wall admit dilution or cooling air to the area between the vanes. The circumferentially spaced apertures 34 provided in the combustor 18 adjacent the leading edges 36 of the vanes 28 have short truncated tubes 38 mounted in each aperture 34 which direct secondary air from the passages 20 and 22 into the combustor and onto the leading edge 36 of the vanes 28.

The primary combustion section 40 is the area within the combustor between the tip of the fuel nozzles 26 and the point of admitting secondary air at the leading edges 36 of the vanes 28 through apertures 34. The forward portion or primary combustion section 40 is fabricated from known techniques so that adequate film cooling of this area is provided by air flowing through the slots 42. The particular combustor illustrated is of the progressive burning type so that primary air is admitted through the larger apertures 44 and smaller aperture 45 in addition to the primary air admitted through passages 19. The progressive type burner forms no part of this invention but suflice it to say that the fuel admitted through nozzles 26 is completely burned by the time it reaches the dilution zone-sufiicient air for a stoichiometeric mixture with a little excess for practical reasons being admitted through passages 19 and apertures 44 and 45.

In operation, compressed air is proportioned by deflector 23 bet-ween passages 19 and the annular passages 20 and 22. The air admitted through passages 19 is mixed with the fuel spray from nozzles 26 and ignited and burned. Additional air from passages 20 and 22 flows through apertures 44 and 45 to completely burn the fuel and through slots 42 to film cool the combustor inner walls. Further downstream, secondary air flows radially through the apertures 34 and is directed by the tubes 38 into the combustor and onto the vane leading edges 36. This flow is substantial and dilutes the combustion gases. In addition, the flow has an impingement cooling elfect on the leading edges 36 of the vanes 28. Air from passages 20 and 22 also flows radially into the hollow vanes from both directions and out the apertures 35 and along their radial walls toward the aft end of the combustor structure 12. This last-mentioned flow provides convection cooling of the vanes 28 and a film of cool air adjacent the outer surface of the vanes 28. Dilution air may also be admitted through the apertures 37 between the vanes 28. The flow of primary or combustion air is indicated by open arrows and the flow of secondary air is indicated by black arrows in FIG. 1. Note the vanes are convection cooled and have their exposed surface film cooled while their most critical areas, that is, their leading edges are impingement cooled by the dilution air flowing through aperture 34. This cooling combination I have found provides adequate cooling of the vanes 28 made of presently available material with approximately 50% of the air flow while the remaining 50% of the compressed air available as primary air for I combustion is adequate to support efiicient combustion and yield satisfactory engine performance. The cooling'of the blade in this manner also aids in presenting a temperature profile at the stator exit which is capable of being withstood by the first stage turbine 46.

Combustors are designed from empirical data and normally the primary combustion zone 40 has a length-toheight ratio of approximately 2.5 whereas the integrated turbine stator-combustor of my invention has a primary combustion zone with a length-to-height ratio on the order of 1.5. Thus my invention provides a turbine-stator-combustor structure capable of efl'icient combustion and with standing the combustion temperatures which is considerably shonter than comparable structures now available.

It should be understood, of course that the foregoing disclosure relates to only a preferred embodiment of the invention and that it is intended to cover all changes and modifications of the example of the invention herein chosen for the purposes of the disclosure, which do not constitute departures from the spirit and scope of the invention.

I claim:

'1. A turbine-stator-combustor for a gas turbine engine or the like comprising in combination: an annular casing, a combustor mounted in said casing and spaced radially therefrom to provide inner and outer annular air passages, a plurality of circumferentially spaced stator vanes mounted at the aft end of said combustor, said combustor having a plurality of circumferentially spaced apertures adjacent the leading edges of said vanes, truncated tubes mounted in said combustor apertures with their maximum height remote from the leading edges of said vanes, said truncated tubes extending into said combustor to direct secondary air onto said leading edges of said vanes to cool them by impingement of the secondary air.

2. The turbine-stator-combustor as defined in claim 1 wherein said stator vanes are hollow, having open inner and outer ends in fluid communication with said annular air passages, respectively, and have a plurality of apertures in their radial walls whereby secondary air is directed into said hollow blades and outwardly along the exposed surfaces of said radial walls to convection cool the said vanes and to provide a film of cool air adjacent .said radial walls.

3. The turbine-stator-combustor as defined in claim 2 wherein said combustor has a plurality of apertures between said vanes.

References Cited UNITED STATES PATENTS 2,625,793 1/ 1953 Mierley et a1 -3955 2,709,894- 6/1955 Onlianoff et a1 60--39.65 3,316,714 5/1967 Smith et al. 60-39.65

SAMUEL FEINBERG, Primary Examiner U.S. Cl. X.R. 

